The present invention relates generally to gas turbine engines, and, more specifically, to turbine flowpaths therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. The gases flow through turbine stages including stator vanes and rotor blades which extract energy from the gases for powering the compressor and producing power for propelling an aircraft in flight for example.
The turbine vanes and blades include airfoils which are bathed in the hot combustion gases during operation, and are therefore typically cooled. Airfoil cooling is effected by bleeding a portion of pressurized air from the compressor and channeling it through the airfoils in various manners of cooling thereof.
Each airfoil may have one or more internal cooling channels which distribute the cooling air therethrough, which is then discharged from the airfoil through various outlet holes formed through the wall thereof.
Since bleeding compressor air for cooling the turbines necessarily decreases overall efficiency of the engine, maximum cooling efficiency of that bled air is desired. The prior art is quite crowded with various forms of internal cooling circuits in turbine vanes, blades, and shrouds, and various forms of discharge holes including film cooling holes and trailing edge cooling holes.
Trailing edge cooling is particularly problematic in view of the relative thinness of the airfoil thereat. An airfoil includes pressure and suction sides which extend from root to tip, and are joined at opposite leading and trailing edges. The pressure and suction sides converge to the trailing edge which may be about 30 mils (0.76 mm) thick or less.
A typical trailing edge cooling design includes a row of axially extending discharge holes spaced apart radially along the longitudinal span of the airfoil. In view of the relative thinness of the airfoil trailing edge, this region of the airfoil is typically solid except for the trailing edge discharge holes typically centered in the wall between the opposite pressure and suction sides. The trailing edge holes extend axially forwardly to a common supply channel in which compressor bleed air is channeled for providing a coolant. The coolant air is discharged through the trailing edge holes for cooling the trailing edge region by internal convection.
The trailing edge cooling holes must necessarily have small diameters to fit within the narrow width of the trailing edge wall. The holes are correspondingly relatively long, with a length-to-diameter ratio up to about 50, for example.
Accordingly, as the coolant flows through the slender trailing edge holes, heat is absorbed from the airfoil by convection. Convection cooling is limited in capability, and therefore the density of the trailing edge cooling holes is typically high for effectively cooling the trailing edge.
As turbine operating temperatures increase, and blade size decreases, the problem of effective trailing edge cooling increases. Additional compressor bleed air may be required for meeting the higher heating demands of the blade resulting in a corresponding reduction in engine efficiency.
Accordingly, it is desired to further increase the cooling efficiency of turbine engine components.